Cooling system for aerospace vehicle components

ABSTRACT

A cooling system for removing heat generated by electrical components onboard aerospace vehicles flows coolant between an evaporator that removes heat from the component, and a condenser within the skin of the vehicle. The skin is formed from facesheets comprising multiple layers of polymer resin reinforced with meshes of single and double wall nanotubes. The nanotubes conduct the heat directionally so as to both distribute the heat over the skin and direct the heat to the outer surface of the skin where the heat can be carried away by air flowing over the skin. The skin may also include conductive carbon foam surrounding the condenser to reduce thermal resistance between the condenser and the facesheets.

PRIORITY CLAIM

The present application is a divisional of U.S. patent application Ser.No. 11/747,467, filed on May 11, 2007 and entitled “Cooling System forAerospace Vehicle Components,” the disclosure of which is incorporatedherein by reference in its entirety.

TECHNICAL FIELD

This disclosure generally relates to systems for removing heat generatedby electrically powered subsystems and components such as electronicdevices onboard an aerospace vehicle, and deals more particularly with acooling system integrated into the structure of the vehicle, such as theskin of an aircraft or spacecraft.

BACKGROUND

The increasing use of avionics and electrically powered subsystemsonboard commercial and military aircraft requires improved thermalmanagement of the heat loads produced by these electrical components.Current solutions to the problem of thermal management may be achievedat the expense of higher costs, possible reduction of overall componentperformance, decreased efficiency and/or increased weight. Effectivemanagement of thermal loads is made more difficult by the trend towardthe use of outer skins formed from non-metallic materials in order toreduce weight. Commonly used structural composites typically have poorthermal conductivity, compared to metal skins. In the case of certainmilitary aircraft, the need to maintain exterior surfaces smooth with aminimum number of penetrations to achieve radio frequency stealth mayfurther reduce the design options for managing thermal loads.

Accordingly, there is a need for an improved cooling system forcontrolling heat loads generated by electrical components onboardaerospace vehicles. The illustrated embodiments of the disclosure areintended to satisfy this need.

SUMMARY

Embodiments of the disclosure transfer heat generated by onboardelectrical components to areas on the aircraft where the heat can bereleased or dissipated. A cooling system for removing the heat may beintegrated into a skin on the aircraft, such as a wing skin. Layers ofresin reinforced with unidirectional carbon nanotubes allow the heat tobe conducted through the thickness of the skin and then spread over theskin surface in order to improve thermal transfer efficiency.

According to one embodiment, a composite skin for aircraft is provided,comprising a thermal distribution system for distributing heat from aheat source laterally through the plane of the skin, and first andsecond facesheets on opposite sides of the thermal distribution system.At least one of the facesheets is thermally conductive for conductingheat from the distribution system to the face of the skin. The facesheetincludes a first layer of material for conducting heat laterally throughthe plane of the facesheet, and a second layer of material contactingthe first layer for conducting heat transversely through the plane ofthe facesheet. The first and second layers are bonded by a thermallyconducted adhesive. The facesheet may include a third layer of materialcontacting the second layer for conducting heat laterally through theplane of the facesheet. The first layer of material may include a meshof carbon nanotubes aligned in one direction and held in a syntheticresin matrix. The second layer of material may include a mesh of doublewall carbon nanotubes aligned in a direction transverse to the plane ofthe first layer and held in a synthetic resin matrix. The thermaldistribution system may include a heat transfer fluid and a serpentineheat exchanger between the first and second facesheets through which theheat transfer fluid may flow.

According to another embodiment, a system is provided for controllingheat generated by electronic components onboard a winged aircraft. Thesystem comprises a cooling system for absorbing heat from an electroniccomponent and transferring the absorbed heat to a wing on the aircraft,and a thermally conductive skin on the wing connected with the coolingsystem for transferring the absorbed heat to air flowing over the wing.The cooling system may include first and second heat exchangersrespectively thermally coupled with the electronic component and theconductive skin. A heat transfer medium flowing between the first andsecond heat exchangers conveys the absorbed heat to the wing. The firstheat exchanger may include an evaporator for vaporizing the heattransfer medium, and the second heat exchanger may include a condenserfor condensing the heat transfer medium.

According to still another embodiment, a thermally conductive facesheetis provided for use in the skin of an aircraft. The facesheet comprisesa first layer of thermally conductive material for conducting heat in afirst direction laterally through the skin, and a second layer ofthermally conductive material contacting the first layer for conductingheat in a second direction transverse to the first direction. Thefacesheet may further comprise a third layer of thermally conductivematerial contacting the second layer for conducting heat in a thirddirection laterally through the skin. The first layer may include a meshof carbon nanotubes aligned in the first direction and held in asynthetic resin matrix. The second layer may include a mesh of doublewall carbon nanotubes aligned in the second direction and held in asynthetic matrix. The third layer may include a synthetic resinreinforced with single wall nanotubes aligned in the third direction.

According to a method embodiment of the disclosure, a method is providedof cooling an electronic component onboard an aircraft. The methodcomprises the steps of: transferring heat from an electronic componentto a coolant; flowing the coolant through at least a section of a wingon the aircraft; and, transferring heat in the coolant to a surface ofthe wing. Heat may be transferred from the electronic component to thecoolant by evaporating the coolant to absorb heat from the electroniccomponent. The flow of coolant may be distributed through the wing bypassing the coolant through a serpentine coil. Heat in the coolant istransferred to a surface of the wing by condensing the coolant torelease the heat contained in the coolant. The heat may be transferredfrom the coolant to the wing surface by spreading the heat across afirst layer of material, conducting the heat from the first layer to asecond layer, and spreading the heat across the second layer. The heatmay be spread across the first and second layers by conducting heatthrough carbon nanotubes.

Other features, benefits and advantages of the disclosed embodimentswill become apparent from the following description of embodiments, whenviewed in accordance with the attached drawings and appended claims.

BRIEF DESCRIPTION OF THE ILLUSTRATIONS

FIG. 1 is a perspective illustration of a wing box having an internalactuator motor, and showing part of the cooling system forming anembodiment of the disclosure.

FIG. 2 is a perspective illustration of the actuator motor removed fromthe wing box and showing the fluid connections to the lower wing skin.

FIG. 3 is a perspective illustration of the actuator motor shown inFIGS. 1 and 2, better illustrating a thermal enclosure clamped aroundthe actuator motor.

FIG. 4 is an exploded, perspective illustration of the thermalenclosure, including an evaporator tube.

FIG. 5 is a schematic illustration of the cooling system loop.

FIG. 6 is a sectional illustration taken along the line 6-6 in FIG. 5.

FIG. 7 is a diagrammatic illustration of an alternate form of a valvethat may be used in the cooling system loop shown in FIG. 5.

FIG. 8 is a sectional illustration of a wing skin forming oneembodiment.

FIG. 9 is a perspective illustration of a facesheet forming part of theskin shown in FIG. 8.

FIG. 10 is a cross sectional illustration of the area designated as “A”in FIG. 9.

FIG. 11 is a plan view illustrating one embodiment of a condenser coil.

FIG. 12 is an illustration similar to FIG. 11 but showing an alternateembodiment of the condenser.

FIG. 13 is a sectional illustration of another embodiment of the wingskin.

FIG. 14 is a sectional illustration of a further embodiment of the wingskin.

FIG. 15 is a sectional illustration of another embodiment of the wingskin.

FIG. 16 is a perspective illustration of a wing skin section employingan alternate embodiment of the condensing coil.

FIG. 17 is an illustration similar to FIG. 16 but showing anotherembodiment of the condensing coil.

DETAILED DESCRIPTION

The embodiments of the disclosure described below relate to a coolingsystem and components thereof for removing and/or distributing heatedgenerated by one or more components or subsystems onboard an aircraft.It should be noted however, that the cooling system may beadvantageously used on a variety of aerospace vehicles, including spacevehicles and satellites. Referring FIGS. 1-5, a wing box 20 for anaircraft includes upper and lower skins 22, 24 (FIG. 1) connectedtogether by longitudinally extending spars 21. A heat generatingcomponent in the form an electric actuator motor 28 is mounted insidethe wing box 20. Motor 28 displaces a shaft 30 through a drive mechanism38 in order to move a wing flap support 32. The heat generated by motor28 within the enclosed space of the wing box 20 may reduce theperformance and/or service life of the motor, and/or create undesirable“hot spots” in the wing that may be detected using infrared sensingtechniques. The actuator motor 28 is merely illustrative of a widevariety of electrical and electronic components and subsystems that maygenerate heat onboard the aircraft.

As best seen in FIGS. 3 and 4, a thermal enclosure 34 includes twohalves 34 a, 34 b that are bolted together so as to clamp enclosure 34around the actuator motor 28. The thermal enclosure 34 is formed of athermally non-conductive, lightweight material such as carbon fiberreinforced epoxy or other composite material. Each of the enclosurehaves 34 a, 34 b includes a pair of cavities 47 for respectivelyreceiving heat sinks 49 formed of thermally conductive material, such asa graphite carbon foam. The heat sinks 49 include curved surface areasfor complementally engaging the cylindrical body of the motor 28, andare secured to an enclosure cover 36.

The enclosure covers 36 are secured to the respective enclosure halves34 a, 34 b using removable fasteners. The enclosure covers 36 are formedof a thermally conductive material such as aluminum and include anintegrally formed outer pocket 45 for receiving a cylindrically shapedevaporator tube 40. Alternatively, the evaporator tubes 40 may be formedintegral with the enclosure covers 36. The carbon foam inserts 49 may befilled with a phase change material to promote the conduction of heatfrom the motor 28 to the evaporator tube 40. A clamp member 44 is boltedto the enclosure cover 36 and tightly clamps the evaporator tube 40 inthe pocket 45. Although not shown in the drawings, a mechanical lock maybe provided between the enclosure 34 and the motor 28 to preventrelative rotation therebetween.

As best seen in FIGS. 1 and 3, a cylindrically shaped compensationchamber 42 is also secured to each of the covers 36. The evaporator tube40 and the compensation chamber 42 are connected by fluid lines 46, 50in the cooling loop shown in FIG. 5. Heat 64 passes from the motor 28through the inserts 49 and the enclosure covers 36 to the walls of theevaporator tube 40. The heat then passes through a cylindrical wick 62and is absorbed by a coolant fluid 58 which may be a two phase, heattransfer fluid such as ammonia, Freon®, water or methanol.

The coolant fluid 58 evaporates as it absorbs the heat 64 to produce ahot vapor that is drawn through line 46 to a later discussed condenser54. Condenser 54 includes condenser coils 52 where the coolant fluid iscooled and condensed into a liquid. During the phase change from a hotvapor to a liquid, heat is output at 66 a or 66 b from the condensercoils 52. The condensed liquid coolant passes through directionalcontrol valve 56 which allows the coolant flow to transfer heat into anupper or lower condenser coil 52. The directional control valve 56 maycomprise, for example, a single coil, piloting solenoid valve.Alternatively, as shown in FIG. 7, the expansion valve 56 may comprise atwo way nanoflap-valve 56 a in which incoming fluid 68 is routed toeither of two exit channels 70, 72 by a micro-flap 56.

The cooled vapor exits the directional control valve 56 and is deliveredby line 50 to the compensation chamber 42 which acts as a buffer-likereservoir. The vapor then passes back into the evaporator 40 wherefurther heat 64 is absorbed, thus completing the cooling cycle.

The condenser coils 52 are positioned inside one or both of the skins22, 24. In the illustrated example, the condenser coils 52 aresandwiched between multiple layers of material which will be describedin more detail below. Referring now to FIGS. 8-11, one embodiment of askin 24 a comprises multiple layers of material which not only provide alightweight, strong skin covering, but function to transfer and releasethe heat generated by the actuator motor 28 over a relatively broad areaof the skin 24 a. It should be noted here that while the skin 24 aillustrated in the present disclosure is a wing skins, the principles ofthe disclosed embodiments may be advantageously used in skins coveringother surfaces of the aircraft, including the fuselage.

The wing skin 24 a broadly comprises inboard and outboard facesheets 74,76, a structural core 78 and one or more condenser coils 52 disposedwithin a layer 80 of thermally conductive material, such as carbon foam.One suitable foam that may be used as the layer 80 is disclosed in U.S.Pat. No. 7,070,755, issued Jul. 4, 2006 to Klett et al. The thermallyconductive foam disclosed in Klett et al normally has a thermalconductivity of at least 40 W/mK, and has a specific thermalconductivity, defined as the thermal conductivity divided by thedensity, of at least about 75 W cm³/m° Kgm. This foam may also have ahigh specific surface area, typically at least about 6,000 m²/m³. Thefoam is characterized by an x-ray diffraction pattern having “doublet”100 and 101 peaks characterized by a relative peak split factor nogreater than about 0.470. The foam is graphitic and exhibitssubstantially isotropic thermal conductivity. The foam comprisessubstantially ellipsoidal pores and the mean pore diameter of such poresis preferably no greater than about 340 microns. Other materials, suchas phase change materials, can be impregnated in the pores in order toimpart beneficial thermal properties to the foam.

The carbon foam layer 80 may be machined to match a desired exteriorcontour of the wing skin 24 a. As best seen in FIG. 11, each of thecondenser coils 52 may include multiple fluid connections 92 thatconnect the coil 52 to the feedline 46 and the expansion valve 56. Inthe embodiment illustrated in FIG. 11, the condenser coil 52 includes aseries of parallel, straight tube sections or legs 52 a that areconnected by curved end sections 94. The longitudinal axes of thestraight tube sections 52 a extend in a fore and aft direction withinthe wing box 20. Through this fore and aft arrangement, bending stressescreated by wing deflection are limited to the curved sections 94 of thecoil 52.

The condenser coils 52 are disposed within slots that may be machined inthe layer 80 of carbon foam and are potted in a thermally conductivepotting compound 82 which may be a graphite filled potting compound, forexample. The potting compound may comprise a paste exhibitingconductivity of at least about 2.5 W/m-K. For example, the paste maycomprise a modified form of a paste available from the HenkelCorporation and identified by the manufacturer as Hysol® EA 9396. Thethermal properties of the paste may be modified by mixing thermallyconductive carbon nanotubes particles in the solution. This paste is alow viscosity, room temperature curing adhesive system with goodstrength properties. A portion of the condenser coils 52 contact theinterior face of the outboard facesheet 76. By placing a portion of thecoils 52 in contact with the facesheet 76 and using a thermallyconductive potting compound 82 to fix the condenser coils 52 in positionwithin the carbon foam 80, the thermal resistance between the condensercoils 52 and the outboard facesheet 76 is reduced. The core 78 may be alightweight structural material such as a honeycomb formed from any of avariety of materials such as aluminum, thermoplastic or NOMEX®.

The construction of the inboard and outboard facesheets 74, 76 may be,but need not be identical. In the illustrated embodiment, the outboardfacesheet 76, which is exposed to free-flowing air over the wing, is asandwiched construction, the details of which are shown in FIGS. 9 and10. Facesheet 76 comprises a central layer of material 88 sandwichedbetween inner and outer layers of material 84, 86 respectively. Layer 84may comprise a synthetic resin matrix such as epoxy reinforced with amesh of aligned single wall nanotubes (SWNT) formed of carbon. In theillustrated example, the SWNTs are magnetically aligned in a desireddirection shown by the arrow 90 so as to conduct heat unidirectionallywithin the layer 84. SWNTs comprise a one atom thick sheet of graphiterolled into a seamless cylinder with a diameter on the order of ananometer. Carbon nanotubes are exceptionally strong and are goodconductors of both electricity and heat.

The mesh of magnetically aligned SWNTs may be formed by forcing asuspension of carbon nanotubes through a fine mesh filter, and thensubjecting the mesh to an electric or magnetic field that aligns thenanotubes in the mesh along one direction. The resulting flat sheets ofmeshed nanotubes, known in the art as “bucky paper”, are then infusedwith a polymer binder such as epoxy, that forms a matrix reinforced bythe carbon nanotube mesh. As shown in FIG. 10, multiple layers 84 a ofthe bucky paper may be used to achieve a desired thickness of thefacesheet 84.

The central layer 88 may comprise a polymer resin infused double wallnanotubes (DWNT), in which double wall nanotubes are arranged in a meshand unidirectionally aligned in the direction of the arrow 85,transverse to the alignment of the nanotubes in the inboard layer 84. Inother words, the DWNTs in layer 88 are aligned in the direction of thethickness of the layer 88. DWNTs comprise multiple layers of graphiterolled in upon themselves to form a double wall tubular shape. In oneform, DWNTs may be arranged in concentric cylinders, forming a singlewall nanotube within a larger single wall nanotube. In another form, asingle sheet of graphite may be rolled in around itself to resemble ascroll. DWNTs possess properties similar to SWNTs but allowfunctionalization of the outer nanotube, i.e. grafting of chemicalfunctions at the surface of the outer nanotube, while maintaining theinner nanotube pristine. DWNTs may offer performance equal or betterthan SWNTs for conduction and emission of electrons, and showsignificantly longer useful lifetimes. DWNTs are commercially availablefrom Tailored Materials Corporation Inc. located in Tucson, Ariz. Thefacesheets 74, 76 may employ a resin matrix exhibiting thermalconductivities up to 5 W/m-K.

As in the case of layers 84, layer 88 may comprise multiple sheets ofmagnetically aligned MWNTs arranged in a mesh and infused with a polymerresin such as epoxy. The layers 84, 86 of SWNTs are bonded to thecentral layer 88 using a layer or film 90 of conductive adhesivematerial that may include carbon nanofibers.

Heat delivered by the condenser coils 52 is conducted both directly tothe outer facesheet 76, and indirectly through the potting compound 82and carbon foam layer 80 to the outer facesheet 76. Thus, the carbonfoam layer 80 assists in transferring heat from almost the entirecircumference of the condenser coils 52 while assisting in spreading theconducted heat over a wider area of the outboard facesheet 76. Thedirectional nature of the SWNT layer 84 further assists in spreading theheat conducted through the carbon foam layer 80. The heat absorbed bythe SWNT layer 84 is then conducted by the central DWNT layer 88 to theSWNT layer 86.

The SWNT layer 86 is magnetically aligned to thermally conduct heatacross the layer 86, thereby dispersing the heat over a wider areabefore it is released into the air flowing over the layer 86. As aresult of the use of the carbon foam layer 80 in combination with thealigned nanotube mesh in layers 84, 86 and 88, the heat released by thecondenser coils 52 is spread over a wider surface area, thereby avoiding“hot spots” that may degrade performance of components and/or act asundesirable radar signatures in military applications.

FIG. 12 illustrates an alternate form of condenser coils 52 a in whichthe tube legs 53 are connected at their ends by fittings 98, rather thanthe bends 94 used in the coil 52 shown in FIG. 11. By using fittings 98to connect adjacent tube legs 53 the total length of a coil 52 may beincreased within a given area of space within the wing box 20.

FIG. 13 illustrates another embodiment of a skin 24 b. In thisembodiment, the condenser coils 52 are bonded directly to the outerfacesheet 76. A conductive carbon foam strip 80 a is fitted around eachof the coils 52. Honeycomb cores 100 are placed between adjacent foamstrips 80 a. The skin construction shown in FIG. 13 may provide areduction in weight compared with the skin 24 a shown in FIG. 8.

Another embodiment of a skin 24 c is shown in FIG. 14. The honeycombcore 78 a is fitted around a portion of each of the condenser coils 52.The condenser coils 52 are positioned within slots in the carbon foamlayer 80 and are bonded directly to the outboard facesheet 76. The skinconstruction shown in FIG. 14 reduces weight as a result of thereduction in the thickness of the carbon foam layer 80, but in somecases may increase the thermal resistance between the condenser coils 52and the outboard facesheet 76.

FIG. 15 illustrates still another embodiment of a skin 24 d in which thecondenser coils 52 are disposed within channels in the carbon foam layer80, and are spaced slightly from the outboard facesheet 76. In thisembodiment, all of the heat removed from the condenser coils 52 mustpass through the carbon foam layer 80.

A further embodiment of the skin 24 e is illustrated in FIG. 16. Thecondenser coils 52 a are alternately attached directly to the inboardand outboard facesheets 74, 76. The coils 52 a each include a flange 52b which provides an increased area of contact with the correspondingfacesheet 74, 76. The coils 52 a are held in potting compound 82 withinchannels formed in a layer 80 of conductive carbon foam. In thisembodiment, heat is conducted to both the inboard and outboardfacesheets 74, 76. The inboard facesheets 74 may face or surround athermal mass such as fuel within a fuel tank that may absorb the heatconducted through the inboard facesheets 74.

Similarly, as illustrated in FIG. 17, in another embodiment of the skin24 f, the condenser coils 52 a may be vertically stacked in pairs andheld within a potting compound 82 that is separated by strips of carbonfoam 80.

Although the embodiments of this disclosure have been described withrespect to certain exemplary embodiments, it is to be understood thatthe specific embodiments are for purposes of illustration and notlimitation, as other variations will occur to those of skill in the art.

What is claimed is:
 1. A system for controlling heat generated by electronic components on-board an aerospace vehicle, comprising: a cooling system, configured to absorb heat from an electronic component on-board the aerospace vehicle, and to transfer the absorbed heat to an exterior surface of the aerospace vehicle; and, a thermally conductive skin, on the exterior surface, configured to transfer the absorbed heat to an environment contacting the thermally conductive skin.
 2. A system in accordance with claim 1, wherein the cooling system includes: a first heat exchanger, thermally coupled with the electronic component; a second heat exchanger, thermally coupled with the conductive skin; and a heat transfer medium, flowing between the first and second heat exchangers, configured to convey the absorbed heat to the thermally conductive skin.
 3. A system in accordance with claim 2, wherein: the first heat exchanger includes an evaporator for vaporizing the heat transfer medium; and the second heat exchanger includes a condenser for condensing the heat transfer medium.
 4. A system in accordance with claim 2, wherein the second heat exchanger includes a serpentine coil through which the heat transfer medium can flow, the serpentine coil including: generally straight legs, disposed within the thermally conductive skin, extending fore and aft relative to a direction of travel of the aerospace vehicle; and sections connecting fore and aft ends of the generally straight legs.
 5. A system in accordance with claim 2, wherein the second heat exchanger includes a serpentine coil including portions in thermal contact with the thermally conductive skin.
 6. A system in accordance with claim 2, further comprising a valve for controlling the flow of the heat transfer medium between the first and second heat exchangers.
 7. A system in accordance with claim 2, wherein: the thermally conductive skin includes inner and outer facesheets; and the second heat exchanger includes a portion in thermal contact with the inner facesheet, the environment contacting the inner facesheet comprising a thermal mass disposed within the aerospace vehicle, whereby at least a portion of the absorbed heat can be transferred to the thermal mass.
 8. A heat exchange system in accordance with claim 7, wherein the thermal mass disposed within the aerospace vehicle comprises a fuel tank.
 9. A system in accordance with claim 1, wherein the thermally conductive skin includes: inner and outer facesheets; and a layer of thermally conductive material, disposed between the inner and outer facesheets, configured to conduct heat from the second heat exchanger to at least one of the inner and outer facesheets.
 10. A system in accordance with claim 9, wherein at least one of the inner and outer facesheets includes: a first layer of material, configured to conduct heat laterally through the plane of the at least one facesheet; and a second layer of material, contacting the first layer of material, configured to conduct heat transversely through the plane of the at least one facesheet.
 11. A system in accordance with claim 10, wherein the at least one facesheet includes a third layer of material, contacting the second layer, configured to conduct heat laterally through the plane of the at least one facesheet.
 12. A system in accordance with claim 10, wherein the first layer of material includes a synthetic resin reinforced with a mesh of carbon nanotubes aligned in one direction within the plane of the at least one facesheet.
 13. A system in accordance with claim 9, wherein at least one of the inner and outer facesheets includes: first and second layers, each layer including a synthetic resin reinforced with unidrectionally aligned single wall carbon nanotubes; and a third layer, disposed between the first and second layers, the third layer including a synthetic resin reinforced with double wall carbon nanotubes aligned in a direction transverse to each of the alignment directions of the single wall carbon nanotubes in the first and second layers.
 14. A system in accordance with claim 9, wherein the thermally conductive material includes thermally conductive carbon foam.
 15. A system in accordance with claim 9, wherein the inner facesheet is in contact with a thermal mass disposed within the aerospace vehicle, whereby at least a portion of the absorbed heat can be transferred to the thermal mass.
 16. A system in accordance with claim 1, wherein the environment contacting the thermally conductive skin includes a thermal mass disposed within the aerospace vehicle.
 17. A heat exchange system for an aerospace vehicle, comprising: a composite skin, having a structural core; a plurality of thermally conductive strips, adjacent the structural core, each strip having a channel integrally formed therein; a plurality of strips of structural core material, interleaved between adjacent strips of the plurality of thermally conductive strips; and inner and outer facesheets, disposed on opposite sides of the structural core, the thermally conductive layer and the strips of structural core material, at least one of the facesheets being thermally conductive; and a thermal distribution system, comprising a heat exchange fluid conduit disposed within the channel in each thermally conductive strip, adapted to distribute heat from a heat source to an environment contacting the composite skin.
 18. A heat exchange system in accordance with claim 17, wherein the heat exchange fluid conduit is in thermal contact with at least one of the inner and outer facesheets.
 19. A heat exchange system in accordance with claim 17, wherein the heat exchange fluid conduit includes a portion in thermal contact with the inner facesheet, the environment contacting the composite skin at the inner facesheet comprising a thermal mass disposed within the aerospace vehicle, whereby at least a portion of the absorbed heat can be transferred to the thermal mass.
 20. A heat exchange system in accordance with claim 19, wherein the thermal mass disposed within the aerospace vehicle comprises a fuel tank. 